The present invention relates to a method for repairing turbine engine components, such as blades and vanes which have airfoils.
Thermal barrier coatings are used on turbine engine components, particularly airfoils, to reduce the metal temperatures and thereby extend the life of the components. Traditionally, thermal barrier coatings are bi-layer systems. A metallic coating called a bond coat, usually MCrAlY with hafnium and silicon or a diffusion aluminide, is applied to a superalloy substrate portion of the component, usually a nickel based or cobalt based superalloy substrate, as the first layer. The bond coat provides oxidation and corrosion protection. An insulating ceramic layer, usually 7 wt % yttria partially stabilized zirconia, is then deposited on the bond coat. The bond coat is typically needed because the underlying nickel based or cobalt based superalloy substrate used in many turbine engine components does not have adequate oxidation resistance.
Recently, a new group of superalloys have been developed that exhibit sufficient overall oxidation resistance. These include but are not limited to a nickel based superalloy containing oxygen active elements such as yttrium and a nickel based superalloy containing reduced sulfur. For turbine engine components formed from these superalloys, the bond coat is eliminated and the ceramic insulating layer is directly deposited on the substrate. Before being approved for service, the components formed from these new superalloys are processed to improve adherence of the ceramic material to the substrate alloy. Components of this type exhibit improved thermal barrier coating spall lives.
Despite these latest developments, turbine hardware components, such as the airfoil portions of blades and vanes, will develop spall and oxidation debris on one or more surfaces. Additionally, the turbine engine components will develop cracks, nicks and dents during use as a result of the extreme environment in which they operate. It becomes necessary to regularly remove the components from a turbine engine after a period of time and refurbish them to remove any spall, oxidation debris, cracks, nicks, and/or dents.
In prior repair techniques, the ceramic insulating layer was removed before any other step. As a result, it was difficult to identify those portions of the turbine engine components where oxidation and/or spall had been located. This has been found to be highly undesirable because portions of the component which require repair are too easily missed.
Accordingly, it is an object of the present invention to provide an improved method for repairing turbine engine components.
It is a further object of the present invention to provide a repair method as above which can be used to repair a wide variety of turbine engine components.
The foregoing objects are attained by the repair method of the present invention.
In accordance with the present invention, a method for repairing a turbine engine component having an insulating ceramic layer broadly comprises removing oxidation debris on the turbine engine component, removing the ceramic layer from the turbine engine component, and thereafter blending exposed portions of the turbine engine component to remove nicks, dents, and/or cracks. The method further comprises removing or replenishing zones depleted in aluminum. Still further, the method comprises restoring a tip of the turbine engine component and thereafter applying tip abrasives to restore cutting ability. Finally, a ceramic coating is applied to the repaired turbine engine component.
Other details of the repair method of the present invention, as well as other objects and advantages attendant thereto, are set forth in the following detailed description.